The present invention relates to a propellant system for a hybrid rocket. In particular, the propellant system utilizes discrete oxidizer packages of oxidizer capable of being stored in the oxidizer section of the hybrid rocket. In embodiments, the present invention relates to an oxidizer package for a propellant system for a rocket, in which the package contains ampoules of compounds that enhance the performance of the propellant system. In particular, the anipoules contain oxidizer. The present invention also relates to a rocket utilizing oxidizer in the form of a plurality of such packages capable of being activated or deployed in single or multiple fashion. In particular embodiments, the present invention relates to oxidizer packages comprising oxidizer and an ancillary compound, in which a matrix of ancillary compound has the ampoules of oxidizer therein. The ancillary compound undergoes exothermic decomposition substantially without consumption of oxygen from the oxidizer.
The following acronyms are used in this application:
Hybrid propulsion systems offer numerous potential advantages over solid and/or liquid propulsion systems. Some potential advantages include high mass fraction, low cost, rapid deployment, reduced storage and transportation restrictions, throttling ability, configurable thrust and mission profiles as well as modifiable plume signatures and low temperature sensitivity. Traditionally, propulsion systems have used liquid oxidizers, which in many cases present handling and storage safety issues.
Solid propulsion systems could provide very high specific impulse by utilizing high performance oxidizers such as AND, HAP, HAN, HNF, NP and the like. Many of these oxidizers offer significant gains on performance, reduced or low toxicity and have desirable exhaust signature characteristics. However, many of these oxidizers suffer from varying degrees and forms of instability, such as photo sensitivity, shock, friction and impact sensitivity, decomposition in the presence of moisture, sensitivity to pH and incompatibility (such as hypergolic reaction) to other propellant materials. A typical example of incompatibility is reaction between HNF and curing agents used in solid propellant binder systems such as HTPB and GAP. Many difficulties have been encountered incorporating the oxidizers into propellant systems, and solutions to particular storage and stability problems often result in compromising the theoretical performance potential. For example, current techniques to synthesize HNF still produce particles with length to diameter ratios of 2:1 to 3:1 with significant variation from the mean. This seriously impacts formulation rheology and can prevent achievement of optimum solids loading, as well as aggravating friction sensitivity during mixing and casting operations.
Storing the oxidizer separately in the motor offers the ability to avoid compatibility issues between these oxidizers and common solid propellant system components and to optimize the physical and environmental requirements of the oxidizer, with the potential to improve the performance of tactical missile systems. A typical example of incompatibility is reaction between HNF and curing agents used in solid propellant binder systems such as HTPB and GAP.
Separate storage of solid and semi-solid oxidizers and expulsion systems have been proposed and demonstrated in the past. Some of the difficulties in these approaches include flow stability, concentration and distribution of oxidizer solids in carrier agents, pressurization and piping system requirements, specialized control valves and system integration. Certain examples of flowable oxidizers can behave as mono-propellants causing flame tracking and catastrophic failure in the delivery and storage systems.
Rocket motors based on gaseous and liquid oxidizers have been successfully demonstrated in the past.
A propellant system for a hybrid rocket using solid or semi-solid oxidizers has now been found, including a method of storage and deployment providing structural and operational benefits. Embodiments of the invention provide an oxidizer package for a propellant system for a rocket in which the oxidizer package comprises oxidizer and optionally ancillary compound, the ancillary compound undergoing exothermic decomposition substantially without consumption of oxygen from the oxidizer. The oxidizer may be in ampoules within the ancillary compound.
Accordingly, one aspect of the present invention provides an oxidizer package for a propellant system for a motor in which the oxidizer is separated from fuel grain, the oxidizer package comprising oxidizer material and an ignition system therefor in a wrapping or sealing material.
In preferred embodiments of the invention, the package is adapted to form a plurality of packages of oxidizer material in a stacked column.
In a further embodiment, the oxidizer packages are adapted to be separated by an inhibitor layer.
In a further embodiment, the oxidizer packages are adapted to be separated by an inhibitor layer.
In other embodiments, the packaging material is formed of material selected from polymer coated aluminum, aluminum coated with a mixture of polymer and magnesium, aluminum coated with a fluoroelastomer and magnesium, pyroalloying metal foil, and a pyrotechnic material consisting of laminates of magnesium foil and fluoropolymer, especially fluoropolymer sheet coated with magnesium or aluminum or alloys of magnesium and aluminum, nitrocellulose or double-based propellant.
In other embodiments, the ancillary compound is selected from the group consisting of a fluoropolymer, a fluoroelastomer, an energetic polymer, a non-energetic polymer, an energetic metal, a crystalline explosive, a polymerized peroxide and a double-based propellant. Preferably, the ancillary compound is a solid propellant formulated for stoichiometric oxygen balance or excess oxygen.
In further embodiments, oxidizer is separated from ancillary compound, said oxidizer being contained in ampoules within the ancillary compound. The oxidizer may be a liquid, a solid or a semi-solid material.
In additional embodiments, there is solid propellant composition formulated for at least stoichiometric oxygen balance.
The ampoules may contain carriers or performance enhancing compounds selected from the group consisting of a fluoropolymer, a fluoroelastomer, an energetic polymer, a non-energetic polymer, an energetic metal, a crystalline explosive, polymerized peroxide and a double-based propellant. The ampoule may be formed from coated aluminum, pyroalloying metal foil, or laminate or sheet of aluminum or magnesium, an energetic polymer, a non-energetic polymer, nitrocellulose and a double-based propellant, especially in which the coated aluminum is selected from polymercoated aluminum, aluminum coated with a mixture of polymer and magnesium, and aluminum coated with a fluoroelastomer and magnesium.
A further aspect of the invention provides a hybrid rocket comprising oxidizer material and fuel grain, the oxidizer material being separated from the fuel grain and being in the form of a single package or plurality of packages of oxidizer material and an ignition system therefor, the oxidizer packages generally conforming to the shape of the rocket.
Another aspect of the invention provides a grid of a pyrotechnic material, especially in which the grid is a matrix, mesh, wool, foamed metal or wires of the pyrotechnic material.
A further aspect of the invention provides a propulsion system for a hybrid rocket comprising oxidizer material in a matrix, mesh, wool, foamed metal or wires of structural or pyrotechnic material, especially in which the matrix, mesh, wool, foamed metal or wires enhance burn rate through thermal conductivity.
Another aspect of the invention provides a.propulsion system for a hybrid rocket comprising a plurality of columns of oxidizer material, the columns being supported by a grid of structural and/or pyrotechnic material, especially in which the pyrotechnic grid material is designed to burn more quickly than the oxidizer material contained therein.